1. Field of the Invention
The present invention relates to a full authority digital engine control system for an aircraft engine.
2. Description of the Related Technology
Engine control systems are commonly referred to using the acronym FADEC, and are intended to control the engine by controlling the intake, for example, of fuel in an engine such as a helicopter turbine, or any other type of engine such as a reactor, a turbojet engine, etc., intended for other types of aircrafts, for example such as an airplane, an airship, etc.
In fact and generally speaking, these systems fall within the field of electronic engine control units (EECUs), and in particular helicopter turbines.
Traditionally, these control systems are connected on the one hand to a mechanism for controlling various operating parameters of the engine, such as fuel intake or air intake, and on the other hand a mechanism forming speed sensors arranged on the engine.
As a general rule, they comprise at least two redundant electronic control units connected to said mechanism forming speed sensors of the engine and a device for switching between said units as a function of their operating status.
Builders and rules require that the suppliers of this type of equipment respect a certain number of requirements in terms of operating safety.
These requirements for example include parameters such as the MTBF (Mean Time Between Failure), breakdown/flight time level, etc.
For some of the aforementioned FADECs, one of the basic requirements is to have at least two redundant electronic control units to control the fuel intake in the engine, which then makes it possible to continue controlling that engine in the event one of the units fails.
Thus, if the electronic control unit currently operating breaks down, the switching device then triggers the switch from the control of that faulty electronic control unit to the redundant electronic control unit.
However, such devices are not fully satisfactory. In fact, the switching between electronic control units only occurs if the faulty electronic control unit declares itself faulty to the switching device or if the aircraft's pilot detects that failure and orders the switch himself.
Thus, FADECs with a double electronic control unit may fail if the currently operating electronic control unit fails without detecting it and/or without indicating it to the switching device of the FADEC.
In that case, the engine control is not transferred to the redundant electronic control unit and the pilot can only perform that switching manually if he identifies the source of the problem quickly enough.
If this is not the case, the engine may run away, or on the contrary there may be a risk of a loss of power or burnout of the engine.
Most FADECs of this type also frequently have an electronic overspeed protection system (EOSPS) for the emergency cut-off of the turbine to prevent deterioration thereof.
In the first scenario considered above, the electronic protection system then causes a cutoff of the turbine without any possibility of restarting the latter part. One can see that this situation may have serious repercussions.
One aim of the invention is therefore to obtain a full authority digital engine control system for an aircraft engine having improved operating safety and availability.